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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2016

Longitudinal Aerodynamic Modeling of the Adaptive Compliant Trailing Edge Flaps on a GIII Airplane and Comparisons to Flight Data

A pair of compliant trailing edge flaps was flown on a modified GIII airplane. Prior to flight test, multiple analysis tools of various levels of complexity were used to predict the aerodynamic effects of the flaps. Vortex lattice, full potential flow, and full Navier-Stokes aerodynamic analysis software programs were used for prediction, in addition to another program that used empirical data. After the flight-test series, lift and pitching moment coefficient increments due to the flaps were estimated from flight data and compared to the results of the predictive tools. The predicted lift increments matched flight data well for all predictive tools for small flap deflections. All tools over-predicted lift increments for large flap deflections. The potential flow and Navier-Stokes programs predicted pitching moment coefficient increments better than the other tools.

aerodynamics↗
NASA NTRS1959

Longitudinal Stability and Control Characteristics at Mach Numbers from 0.70 to 2.22 of a Triangular Wing Configuration Equipped with a Canard Control, a Trailing-Edge-Flap Control or a Cambered Forebody

Results of an investigation to determine the static longitudinal stability and control characteristics of an aspect-ratio-2 triangular wing and body configuration equipped with either a canard control, a trailing-edge-flap control, or a cambered forebody are presented without analysis for Mach numbers from 0.70 to 2.22. The canard surface had a triangular plan form and a ratio of exposed area to total wing area of 7.8 percent. The hinge line of the canard was in the extended wing chord plane, 0.83 wing mean aerodynamic chord ahead of the reference center of moments. The trailing-edge controls were constant-chord full-span flaps with exposed area equal to 10.7 percent of the total wing area. The cambered body was a modified Sears-Haack body with camber only ahead of the wing apex. Data are presented for various canard and flap deflections at angles of attack ranging from -6 deg to +18 deg.

Boyd, John W.↗
NASA NTRS1958

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 deg the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 deg flared afterbody. The tests were made through an angle-of-attack range of -2 deg to 20 deg at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge -flap configuration. The flared -after body configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift drag ratio.

Robinson, Ross B↗
NASA NTRS1958

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 degrees; the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 degrees flared afterbody. The tests were made through an angle-of-attack range of -2 degrees to 20 degrees at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared-afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge-flap configuration. The flared-afterbody configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift-drag ratio.

Bernot, P. T.↗
NASA NTRS1976

The Effects of Blowing Over Various Trailing-edge Flaps on an NACA 0006 Airfoil Section, Comparisons with Various Types of Flaps on other Airfoil Sections, and an Analysis of Flow and Power Relationships for Blowing Systems

The results are presented of a two-dimensional investigation conducted to determine the effect of blowing over various types of trailing-edge flaps on a wing having the NACA 0006 airfoil section and a drooped-nose flap. The position and profile of the trailing-edge flap, the nozzle height, and the location of the flap with respect to the nozzle were found to be important variables. Data from many investigations were used to make an evaluation of the effects of blowing on lift. An analysis was made of flow and power relationships for blowing systems.

Dods, J. B., Jr.↗
NASA NTRS1958

Area-Suction Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying area-suction boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effect of boundary-layer control in the handling qualities and operation of the airplane, particularly during landing. The wind-tunnel and flight tests indicated that area suction applied to the trailing-edge flaps produced significant increases in flap lift increment. Although the flap boundary-layer control reduced the stall speed only slightly, a reduction in minimum comfortable approach speed of about 12 knots was obtained.

Cook, Woodrow L↗
NASA NTRS1999

Simulation of Acoustic Scattering from a Trailing Edge

Three model problems were examined to assess the difficulties involved in using a hybrid scheme coupling flow computation with the the Ffowcs Williams and Hawkings equation to predict noise generated by vortices passing over a sharp edge. The results indicate that the Ffowcs Williams and Hawkings equation correctly propagates the acoustic signals when provided with accurate flow information on the integration surface. The most difficult of the model problems investigated inviscid flow over a two-dimensional thin NACA airfoil with a blunt-body vortex generator positioned at 98 percent chord. Vortices rolled up downstream of the blunt body. The shed vortices possessed similarities to large coherent eddies in boundary layers. They interacted and occasionally paired as they convected past the sharp trailing edge of the airfoil. The calculations showed acoustic waves emanating from the airfoil trailing edge. Acoustic directivity and Mach number scaling are shown.

Singer, Bart A.↗
NASA NTRS2000

Stator Loading Measurements Behind a Fan With Trailing Edge Blowing

The problem of aircraft noise pollution around airports has become increasingly important as those areas have become more densely populated. Currently, the removal of older noisier aircraft from operation is reducing noise levels around airports; however, with air traffic projected to increase by about 5% over the next decade the number of commercial aircraft operating in the world is expected to be about 17,700 by the year 2007. To keep noise levels around airports from increasing as a result of traffic increases, it is important to investigate new methods of noise reduction. The objective of this work is to provide a better understanding of the effects that trailing edge blowing has on stator unsteady loading. This is done by presenting flowfield and stator loading data from experiments conducted with and without trailing edge blowing.

Waitz, Ian A.↗
NASA NTRS1999

Implementation of a Trailing-Edge Flap Analysis Model in the NASA Langley CAMRAD.MOD1/Hires Program

Continual advances in rotorcraft performance, vibration and acoustic characteristics are being sought by rotary-wing vehicle manufacturers to improve efficiency, handling qualities and community noise acceptance of their products. The rotor system aerodynamic and dynamic behavior are among the key factors which must be addressed to meet the desired goals. Rotor aerodynamicists study how airload redistribution impacts performance and noise, and seek ways to achieve better airload distribution through changes in local aerodynamic response characteristics. One method currently receiving attention is the use of trailing-edge flaps mounted on the rotor blades to provide direct control of a portion of the spanwise lift characteristics. The following work describes the incorporation of a trailing-edge flap model in the CAMRAD.Mod1/FHUS comprehensive rotorcraft analysis code. The CAM-RAD.Mod1/HIRES analysis consists of three separate executable codes. These include the comprehensive trim analysis, CAMRAD.Mod1, the Indicial Post-Processor, IPP, for high resolution airloads, and AIRFOIL, which produces the rotor airfoil tables from input airfoil section characteristics. The modifications made to these components permitting analysis of flapped rotor configurations are documented herein along with user instructions detailing the new input variables and operational notes.

Charles, Bruce↗
NASA NTRS2017

Transonic Drag Reduction Through Trailing-Edge Blowing on the FAST-MAC Circulation Control Model

A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.

Chan, David T.↗
NASA NTRS1984

Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections

Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.

Brooks, T. F.↗
NASA NTRS1989

The role of shock induced trailing-edge separation in limit cycle oscillations

The potential role of shock induced trailing edge separation (SITES) in limit cycle oscillations (LCO) was established. It was shown that the flip-flop characteristics of transition to and from SITES as well as its hysteresis could couple with wing modes with torsional motion and low damping. This connection led to the formulation of a very simple nonlinear math model using the linear equations of motion with a nonlinear step forcing function with hysteresis. A finite difference solution with time was developed and calculations were made for the F-111 TACT were used to determine the step forcing function due to SITES transition. Since no data were available for the hysteresis, a parameter study was conducted allowing the hysteresis effect to vary. Very small hysteresis effects, which were within expected bounds, were required to obtain reasonable response levels that essentially agreed with flight test results. Also in agreement with wind tunnel tests, LCO calculations for the 1/6 scale F-111 model showed that the model should have not experienced LCO.

Cunningham, Atlee M., Jr.↗
NASA NTRS1992

SIMS chemical analysis of extended impact features from the trailing edge portion of experiment AO187-2

One hundred capture cells from the trailing edge, which had lost their cover foils during flight, were optically scanned for extended impact features caused by high velocity projectiles impinging on the cells while the foils were still intact. Of the 53 candidates, 24 impacts were analyzed by secondary ion mass spectroscopy for the chemical composition of the deposits. Projectile material was found in all impacts, and at least 75 percent of them appear to be caused by interplanetary dust particles. Elemental ratios are fractionated, with refractory elements enriched in the impacts relative to interplanetary dust particles collected in the stratosphere. Although this could be due to systematic differences in the compositions, a more likely explanation is volatility fractionation during the impact process.

Amari, Sachiko↗
NASA NTRS2016

Flat Plate Wake Velocity Statistics Obtained With Circular And Elliptic Trailing Edges

The near wake of a flat plate with circular and elliptic trailing edges is investigated with data from direct numerical simulations. The plate length and thickness are the same in both cases. The separating boundary layers are turbulent and statistically identical. Therefore the wake is symmetric in the two cases. The emphasis in this study is on a comparison of the wake-distributions of velocity components, normal intensity and fluctuating shear stress obtained in the two cases.

CFD↗